Future space programs are planned to send satellites into highly elliptical orbits such as capture orbits, possibly to deploy a lander immediately after capture or after some time spent in this capture orbit. In such orbits, given certain geometrical conditions, eclipses of long duration may occur. During these eclipses, the satellite is deprived of solar energy. The power supply and thermal regulation subsystems of the satellite must then be sufficiently rated to be able to operate for the duration of these eclipses.
Insertion trajectories leading to long eclipses are not generally used because they lead to an increase in the weight of the power supply and thermal regulation subsystems. However, in the situation where landing is initiated from the capture orbit, this can prevent a landing under good conditions (for example in sunlight) at a given latitude. This problem is all the more likely to arise when landing is effected not immediately but after some time spent in the capture orbit to avoid unfavourable landing conditions (for example periods of sandstorms on Mars or conjunctions between the planet and Earth).
In a context subject to the constraint of no modification of the shape of the orbit (for example so that orbit exit manoeuvres remain compact), the solution of changing to a circular orbit at low altitude cannot be envisaged. Nevertheless, one possible solution consists in increasing the capacity and therefore the size and the weight of the batteries onboard the spacecraft. This solution is not entirely satisfactory because the batteries then represent a high mass competing with the mass of the payload and/or occupying too much space onboard the craft. Moreover, this solution may not be sufficient for particularly long eclipses.